Nutation damper

ABSTRACT

A viscous nutation damper for a spin stabilized satellite said damper comprising a sealed arcuate tube disposed near the periphery of the spinning satellite. Said tube is oriented generally parallel to the spin axis and is concave toward the spin axis. A viscous fluid and spherical steel ball are sealed within the tube.

United States Patent [191 Williams, deceased NUTATION DAMPER Donald D.Williams, deceased, late of Inglewood, Calif. by Gloria M. Williams,widow Inventor:

Hughes Aircraft Company, Culver City, Calif.

June 29, 1970 Assignee:

Filed:

Appl. No.: $6,034

Related U.S. Application Data Continuation of Ser. No. 534,251, March14, 1966, abandoned, which is a continuation of Ser. No. 22,733, April18, 1960, abandoned.

us. Cl. ..244/15 A, 74/55, 74/574 Int. Cl. ..B64c 17/00, G01C 19/04Field of Search ..74/5.5, 573, 574;

[56] References Cited UNITED STATES PATENTS 2,584,222 2/1952 OConner..74/5.5 2,966,074 12/ 1960 Rumsey 3,034,745 5/1962 Stewart ..244/155Primary ExaminerMilton Buchler Assistant ExaminerStephen G. KuninAttorney-James K. Haskell [57] ABSTRACT A viscous nutation damper for aspin stabilized satellite said damper comprising a sealed arcuate tubedisposed near the periphery of the spinning satellite. Said tube isoriented generally parallel to the spin axis and is concave toward thespin axis. A viscous fluid and spherical steel ball are sealed withinthe tube.

7 Claims, 11 Drawing Figures Asia!!! [PATENTEDHAY 1:915 0, 57

SHEET 2 [1F 4 PATENTEU HAY H973 r t T 7 PATENTEUW Hm SHEET 0F 4 .Jra 11.

NUTATION DAMPER This is a continuation of prior co-pending applicationSer. No. 534,251, filed Mar. 14, 1966, now abandoned which is acontinuation application Ser. No. 22,733, filed Apr. 18, 1960 nowabandoned.

The present invention relates to space vehicles such as satellites and,more particularly, to a method and apparatus for controlling thevelocity, and the orientation of the spin axis, ofa spin-stabilizedspace vehicle.

Man-made satellites placed into orbit around the earth are oftenprovided with equipment requiring the satellites to be accurately placedinto specific orbits and to be oriented in a predetermined manner. Forexample,ja satellite for use as a radio communication relay may have tobe accurately placed into a west-toeast circular orbit in the plane ofthe earths equator and having a period of 24 hours. Such an orbit isdesirable because the satellite hovers above a single point on theearth, inasmuch as both the satellite and the earth have the sameangular velocity. A satellite that hovers I above a single point on theearth must be accurately placed into an orbit having a 22,750 nauticalmile radius from the center of the earth and the satellite must travelaround the orbit with a linear velocity of 10,090 feet per second.

An active communication satellite, that is, one which is equipped toreceive and retransmit radio waves, is provided with an antenna and maybe provided with solar cells as a source of power, both of which requireaccurate orientation of the satellite to be useful. To increase theantenna gain, the satellite should be spinning about the antenna axisand this axis should be parallel to the earths axis. In this way, theantenna radiation pattern may be omnidirectional about'the antenna axisbut have a narrow beam width about a plane extending through the centerof the satellite perpendicular to the antenna axis. This system providesincreased antenna gain, 8 decibels, for example, in the direction of theearth and permits the use of a radio transmitter having a relatively lowpower output, thus reducing size and weight requirements. Further, solarcells. need be placed only on those surfaces of the satellite thatintercept maximum light from the sun.

It may be found that the satellite drifts relative to a precisestationary orbit and requires correction over a period of time. Thepossible drift due to errors in the velocity of the satellite has beendetermined to be 39.4 per year per foot per second of velocity error.

To keep the cost of the satellite and the launching rocket as low aspossible, the satellite and its apparatus for orientation andstabilization should be simple in operation, light in weight, and smallin size.

Accordingly, it is an object of the present invention to provideapparatus for orienting a spin-stabilized vehicle, such as a satellite,and also for controlling the velocity of the satellite for establishinga desired orbit.

Anotherobject of the invention is the provision of apparatus forcorrecting the orbit of a spin-stabilized vehicle, such as a satellite.

Another object of the invention is the provision of apparatus forprecessing the spin axis of a spin-stabilized body from a first positionto a second position perpendicular to the first position.

Another object of the invention is the provision of means for dampingthe nutation of a spin-stabilized body.

Yet another object of the present invention is to provide increasedantenna gain in a spinning vehicle, such as a space satellite.

A further object of the invention is the provision of optimum solar cellillumination in an orbiting space satellite.

Still another object of the present invention is to provide apparatusfor orienting a spin-stabilized satellite that is simple in form,reliable in operation, small in size, light in weight, inexpensive, andlow in power consumption.

In accordance with these and other objects of the invention, an orbitingsatellite having a radio antenna and solar cells is oriented withrespect to the earth and the sum to optimize the satellite antenna gainand the solar cell illumination. The satellite enters its orbit aroundthe earth with its spin-axis initially oriented perpendicular to theearths axis. The desired orientation of the satellite spin axis isparallel to the earths axis and means is provided for precessing thespin axis of the satellite to the desired orientation by applying areactive force to the satellite in a proper plane. Nutation of thesatellite is damped by a viscous damper provided therein. Means isprovided for sensing the orientation of the satellite relative to theearth, the sun, or both.

Deviations from the correct orbital period, eccentricity, and phase aredetermined from observations made from the earth. Means is provided tocorrect orbital deviations by applying a reactive force to the satellitewith a predetermined amount of force and in the proper direction.

The following specification and the accompanying drawings respectivelydescribe and illustrate an exemplification of the present invention.Consideration of the specification and the drawings will provide acomplete understanding of the invention, including the novel featuresand objects thereof. Like reference characters are used to designatelike parts through the figures of the drawings.

FIG. 1 is an elevation of an exemplary embodiment of a satellitelaunching vehicle in accordance with the present invention showing therelationship of the propulsion rockets to the satellite;

FIG. 2 is an enlarged elevation of a portion of the launching vehicle ofFIG. 1 showing the relationship of fourth, fifth, and sixth stagerockets to the satellite and enclosing heat shield;

FIG. 3 is a plan of the satellite showing an antenna provided thereon ina folded position;

FIG. 4 is a perspective of the satellite illustrating the antenna in itserected position and the solar sensing arrangement;

FIG. 5 is a sectional elevation of a portion of the satelliteillustrating the arrangement of an attitude and velocity control systemincluding compressed gas tanks, valves and nozzles and showing anutation damp- FIG. 9 is a diagram illustrating the change in satelliteorientation resulting from the reactive force depicted as being appliedin FIG. 8; I

FIG. 10 is a diagram illustrating further reactive force being appliedto the satellite; and

FIG. 11 is a diagram depicting the satellite after orientation byprecession has aligned the spin axis of the satellite with the earthsaxis.

Although the present invention does not embrace a vehicle for conveyinga satellite to an orbit, a brief description ofa representative rocketvehicle useful for this purpose is included. Reference is hereby made tothe book Rocket Encyclopedia Illustrated, edited by J. W. Herrick and E.Burgess, Aero Publishers, Inc., Los Angeles, California, 1959, and tothe bibliography therein for details of rocket vehicles and definitionsof terms.

There is shown in FIG. 1 an exemplary embodiment of a rocket vehicle,indicated generally by the numeral 25, for accurately launching a spacevehicle, such as a satellite 26, into a predetermined orbit around theearth. In the present example, the orbit into which the satellite 26 isto be placed is a so-called stationary or synchronous orbit which is acircular west-to-east equatorial orbit having a period of 24 hours. Thisorbit has a radius of 22,750 nautical miles from the center of the earthand the satellite 26 travels at a velocity of 10,090 feet per second(fps) in the direction of the earths rotation and appears to stand stillor hover over one point on the earths surface. The satellite 26 may thenbe used, for example, to relay radio communications over long distances.

The primary or booster portion of the rocket vehicle is a multistagerocket power plant designated by the numeral 27 in FIG. 1 and comprisingfirst, second, third, and fourth stage rockets 31, 32, 33, and 34,respectively, arranged in tandem. The first, second, and third stagerockets 31, 32, and 33 are provided with guidance elements, such as jetvanes (not shown) which are disposed in the jet stream from the rocketexhaust nozzles to deflect the jet and thus obtain a turning force tocontrol the direction and attitude of the vehicle 25. Such anarrangement is shown on page 248 of the above-referenced RocketEncyclopedia. The fourth stage rocket 34 is provided with means, such asspin nozzles or spin rockets (not shown) to rotate the fourth stagerocket 34 about its longitudinal axis for stabilization and sucharrangements are shown and described on pages 456 and 457 of the RocketEncyclopedia. The ignition of the rocket engines, jettisoning of burnedout rockets, guidance, and spin stabilization of the multistage rocketpower plant 27 is automatically controlled by a guidance unit 28, whichmay be one of several types of programming and control systems that arewell known in the art of rocketry.

If desired, the multistage rocket power plant 27 may be, for example,the Scout rocket developed for the National Aeronautics and Space Agencyby Vought Astronautics Division of Chance Vought Aircraft, Inc., Dallas,Texas. Reference is made to the publication entitled Space ResearchVehicle Systems Developed from NASA SCOUT, publication number E9R-l2402, dated August 1959 and published by Chance Vought Aircraft, Inc.,for details of the Scout rocket. In particular, on page 23 of the ChanceVought publication, there is a description of a guidance systemdeveloped by the Minneapolis-Honeywell Regulator Co. which may be usedas the guidance unit 28 in the present rocket vehicle 25.

The multistage rocket power plant 27 propels a gross payload 29 to apoint near the perigee of the transfer ellipse; that is, near the lowestpoint of the elliptical trajectory from the outer atmosphere of theearth to the desired orbit. The gross payload 29, best seen in FIGS. 1and 2, comprises a fifth stage rocket 35, a sixth stage rocket 36 andthe satellite 26. The gross payload 29 and the fourth stage rocket 34are covered by a nose shell or cylindrical heat shield 37 until thelaunching vehicle 25 leaves the earths atmosphere. The heat shield 37 isthen automatically separated from the gross payload 29 and the fourthstage rocket 34. The construction and method of separation of a typicalnose shell is illustrated on page 6 of the Chance Vought publication.

The fifth stage solid-propellant rocket 35, which is secured in tandemto the fourth stage rocket 34, provides the additional thrust requiredfor the satellite 26 to reach the perigee of the transfer ellipse. Thefifth stage rocket 35 is provided with means, such as an electricallyfused annular explosive charge 38 (FIG. 2) which encircles the case ofthe fifth stage rocket 35 for reducing the thrust to zero. Whendetonated, the explosive charge 38 ruptures the combustion chamber ofthe fifth stage rocket 35.

The sixth stage rocket 36 provides the additional thrust to establishthe satellite 26 in the circular orbit from the apogee or highest pointof the transfer ellipse. The sixth stage rocket 36 is secured to thesatellite 26 on the side opposite the side to which the fifth stagerocket 35 is secured, and is oriented in the opposite direction. Thatis, the sixth stage rocket 36 is oriented so that the rocket exhaustnozzle extends out from the satellite 26 in the opposite direction fromthat in which the exhaust nozzle of the fifth stage rocket 35 extendsout from the satellite 26. As will be fully apparent hereinafter, thisis necessary so that the thrust of the sixth stage rocket 36 will beapplied in the correct direction.

As may be seen in FIGS. 3, 4 and 5, the satellite 26 is ofa cylindricalconfiguration and may have a diameter of 29 inches, a thickness of 12/2inches, and a weight on the order of 25 pounds, for example. An antenna50 is secured to one of the plane faces of the satellite 26. After thesatellite 26 has reached its predetermined orbit, the antenna 50 extendsoutwardly from the satellite along an axis extending through the centerthereof coaxial with the longitudinal axis of the rocket vehicle 25.This is the spin axis and the antenna axis of the satellite 26. Theantenna 50 provides a radiation pattern that is substantiallyomnidirectional about the antenna axis, but has a narrow beam widthabout a plane extending through the center of the satellite 26perpendicular to the axis of the antenna 50. The cylindrical surface ofthe satellite 26 parallel to the axis of the antenna 50 is provided withsolar cells, indicated by the numeral 52, for converting sunlight intoelectrical enery- As shown in FIG. 2, the fifth and sixth stage rockets35 and 36 are secured to the plane faces of the satellite 26 and areadapted to be separated from the satellite when desired. The antenna 50,which is a dielectric loaded coaxial transmission line havingcircumferential slots spaced a half wavelength apart, is pivoted at thesurface of the satellite 26. Thus, initially the antenna 50 is foldedagainst the surface of the satellite 26 beneath the fifth stage rocket35 (see FIGS. 2 and 3). The antenna 51) is spring loaded at the pivotpoint for erection thereof after separation of the fifth stage rocket 35from the satellite 26 (FIG. 4).

The satellite 26 is provided with a source of reactive power such as twotanks 53 of gas or condensed liquid and its associated vapor underpressure, for example, compressed nitrogen gas at a pressure of 3,000pounds per square inch. In the present embodiment, the tanks 53 are of atoroidal configuration and are disposed immediately within thecylindrical surface of the satellite 26. A velocity control valve 54 andan attitude control valve 55 are connected to the tanks 53 (best seen inFIG. 5). The valves 54 and 55 are quick acting, lowleakagesolenoid-controlled valves such as the type AF56C-37A manufactured bythe Eckel Valve Company of San Fernando, California. The velocitycontrol valve 54 is connected to a nozzle 56 disposed at the peripheryof the satellite 26 and oriented to provide a jet of nitrogen gasdirected radially outward from the center of the satellite 26 along aline perpendicular to the spin axis passing through the center ofgravity of the satellite 26. The attitude control valve 55 is connectedto a nozzle 57 disposed near the periphery of the satellite 26 andoriented to provide a jet of nitrogen gas directed outward from thesatellite 26 along a line parallel to the spin axis of the satellite 26.

Means is provided for absorbing nutation energy such as a nutationdamper 58 comprising a sealed arcuate tube 60 which may be four inchesin length disposed near the periphery of the satellite 26 (FIG. 5) andoriented generally parallel to the spin axis, the curvature of the tube60 being concave toward the spin axis. The tube 60 is filled with aviscous fluid 61 such as oil or silicone fluid and contains a sphericalsteel ball 62, 54 inch in diameter, for example. The radius of curvatureof the tube 60 may be made equal to its distance from the spin axisdivided by L4 and multiplied by where I, is the moment of inertia of thesatellite 26 around any axis normal to the spin axis passing through thecenter of the satellite 26 and 1 is the moment ofinertia of thesatellite 26 about the spin axis. The factor of 1.4 is necessary becausethe ball 62 rolls rather than slides in the tube 60. This choice ofradius makes the damper resonant at the correct frequency for any spinspeed of the vehicle.

The satellite 26 is also provided with means for sensing its orientationwith respect to the sun. In FIG. 4 there is shown a slit 70 provided inthe cylindrical surface of the satellite 26 providing a fan-shapedangular field of view that extends parallel to the antenna axis. Asingle orientation sensing solar cell 71 is disposed within thesatellite 26 adjacent the slit 70 so that when the sun is within thefield of view, which may be 140, for example, the solar cell 71 isenergized to develop a potential at its output terminals. A second slit72 is provided on thesame face of the satellite 26 and provides a secondfan-shaped angular field of view that intersects the field of view ofthe first slit '70 at an angle which may be 35, for example. Similarly,a second orientation sensing solar cell 73 disposed within the satellite26 develops a potential when the sun is in the field of view of thesecond slit 72.

As may be seen in FIG. 6, the satellite 26 is provided with a radiotransmitter and receiver that is electrically connected to the antenna50 as indicated. A source of potential, such as a storage battery 91,applies electrical power to the radio transmitter and receiver 90. Thesolar cells 52 disposed on the outer surface of the satellite 26 areconnected through rectifiers or charging diodes 92 to the battery 91 tomaintain it in a charged condition. Approximately 2200 solar cells 52may be provided and interconnected in banks. The cells 52 in each bankare connected in a series-parallel arrangement, and although .there maybe different numbers of cells 52 in each bank, the number of cells inseries in each series-parallel arrangement is identical to provide theproper voltage for the battery 91. The charging diodes 92 arenonconductive during periods that the voltage developed by any bank ofcells 52 drops below that of the battery 91.

A radio control circuit 93 is also connected to the radio transmitterand receiver 90 and to the battery 91. The radio control circuit 93 isresponsive to control signals received by the radio transmitter andreceiver 90 for applying a potential from the battery 91 to the variouselectrically controlled devices associated with the satellite 26 andwith the fifth and sixth stage rockets 35 and 36. The particular radioremote control system utilized may be one of several systems well knownin the art, for example, one utilizing subcarrier signals transmitted ona carrier wave. The radio control circuit 93 may, for example, include anumber of filters for separating the various subcarrier signals andactuating relays in response thereto, as is well known in the art.

The explosive charge 38 for rupturing the combustion chamber of thefifth stage rocket 35 is connected to the output of the radio controlcircuit 93. The igniter for the sixth stage rocket 36 is also connectedto the output of the radio control circuit 93. The velocity controlvalve 54 and the attitude control valve 55 are each individuallyconnected to the output of the radio control circuit 93.

The first and second orientation sensing solar cells 71 and 73 arerespectively connected to first and second orientation signaloscillators 95 and 96. The orientation sensing solar cells 71 and 73supply electrical power for the oscillators 95 and 96 so thatorientation signals are developed when the solar cells 71 and 73 areilluminated by the sun. The output signals from the oscillators 95 and96 are applied to the radio transmitter and receiver 90 for transmissionto a satellite control point (not shown). Thus, if the sun passesthrough the field of view of the second slit 72, a signal will bedeveloped by the second oscillator 96; and if the sun passes through thefield of view of the first slit 70, a signal will be developed by thefirst oscillator 95.

In the construction of the satellite 26, the mass of the unitsassociated with the satellite 26 is accurately determined and theequipment is distributed within the satel-.

lite 26 so that the center of gravity is made to coincide with thecenter of the satellite 26; and the axis of maximum moment of inertia ismade to coincide with the antenna axis.

Referring now to FIG. 7, the earth is represented by the circledesignated 100 and is rotating in the counterclockwise direction,indicated by the arrow 101, around an axis 102 represented as going intothe drawing through the north pole. The rocket vehicle 25 is fired froma point 103 on the equator of the earth 100, which point may be, forexample, Jarvis Island in the Pacific Ocean that is located at 23minutes south latitude and 160 west longitude. Prior to firing, thebattery 91 is completely charged, the nitrogen tanks 53 are pressurized,and the radio transmitter and receiver 90 and the radio control circuit93 are placed in operation. After firing, the first four rockets 3134 ofthe propulsion system 27 are automatically fired in sequence and guidedby the guidance unit 28. After burnout, each empty rocket case isjettisoned, as is well known in the art.

After the rocket vehicle 25 has attained considerable altitude, therocket vehicle is automatically turned in an easterly direction tocoincide with the direction of rotation of the earth 100, by theguidance unit 28. Prior to ignition of the fourth stage rocket 34, andseparation of this rocket from the third stage rocket, the fourth stagerocket 34 has a spin about its longitudinal axis imparted to it at arate of 2.7 revolutions per second (rps) to provide spin stabilization.The multistage rocket power plant 27 propels the gross payload 29 to apoint above the earth 100 near the lowest point or perigee 104 of anelliptical trajectory or transfer ellipse 105.

The fifth stage rocket 35 provides the additional thrust necessary tocause the satellite 26 to reach the perigee 104 and to traverse thetransfer ellipse 105. Because the velocity of the satellite 26 at theperigee 104 is quite critical in order to achieve the correct apogee,the velocity is determined from the earth 100 in a well known manner,such as by means ofa radio interferometer, a tracking antenna, or bydoppler frequency shift measurement techniques, for example. Thesemeasurements may make use of the radio transmitter and receiver 90 inthe satellite 26 as a radio repeater.

When the correct velocity has been attained, a radio control signal istransmitted from the earth 100 to the satellite 26 to fire the explosivecharge 38 and rupture the combustion chamber of the fifth stage rocket35 to reduce the thrust provided thereby to zero. The fifth stage rocket35 is then separated from the satellite 26.

A first arrow 106 (FIG. 7) indicates the orientation of the sixth stagerocket 36 and the satellite 26 at the perigee 104 of the transferellipse 105 with the arrow pointing in the direction of thrust of thefifth stage rocket 35. Inasmuch as the satellite 26 is spin-stabilizedabout the rocket axis by the spin imparted by the spin nozzles or spinrockets (not shown) on the fourth stage rocket 34, the satellite 26 andthe sixth stage rocket 36 maintains its attitude in space (arrows 107,108, and 110) as it traverses the transfer ellipse 105 to the other sideof the earth 100. As described hereinbefore the sixth stage rocket 36 issecured to the satellite 26 with an orientation such that it appliesthrust in the direction opposite to that of the thrust of the fifthstage rocket 35.

At the highest point or apogee 109 of the transfer ellipse 105, thesatellite 26 has attained the altitude of the desired circular 24-hourorbit 112, but is traveling at only 5200 fps, which is less than thatrequired for establishment of the satellite 26 in the orbit 112. A radiocontrol signal is transmitted to the satellite 26 to cause ignition ofthe sixth stage rocket 36 to provide the additional velocity of 4890 fpsto establish the satellite 26 into the circular orbit 1 12. At theapogee 109 on the other side of the earth 100 from the firing point 103,the direction of thrust of the sixth stage rocket 36 is such as to causethe satellite 26 to enter the circular orbit 112 due to the fact thatthe sixth stage rocket 36 has maintained its attitude in space whiletraversing the transfer ellipse 105 and due to the fact that theorientation of the sixth stage rocket 36 is opposite to that of thefifth stage rocket 35.

As the satellite 26 traverses its orbit 112, it is spinning about itsspin axis with an angular velocity of 2.7 rps imparted to it by thefourth stage rocket 34. However, the spin axis of the satellite 26 isperpendicular to the earth s axis 102, and thus the antenna does notradiate efficiently toward the earth 100. Accordingly, the satellite 26is reoriented by precessing its spin axis through 90. Precession isaccomplished by the reactive force produced by a jet of nitrogen gasfrom the attitude control nozzle 57 which applies thrust parallel to thespin axis near the periphery of the satellite 26. The jet of nitrogengas is controlled by the attitude control valve to produce a net torquearound an axis perpendicular to the spin axis of the satellite 26. Byperiodically pulsing the jet to be on during only a predeterminedportion of the spin cycle of the satellite 26, the torque is applied inthe correct plane to precess the spin axis through 90 until it isparallel to the earths axis 102. The attitude control valve 55 may beactuated during only approximately 60 of the spin cycle of the satellite26, for example.

The attitude control valve 55 is pulsed by radio control from the earth100. The correct phase of the spin cycle to actuate the valve 55 isdetermined from the earth 100 by means of the first orientation sensingsolar cell 71 adjacent the slit in the satellite 26 and its associatedoscillator which modulates the radio transmission from the satellite 26.By correcting for the twoway propagation delay, the jet is turned onduring the correct portion of the spin angle of the satellite 26 tocause precession in the proper direction. This action is indicated inFIGS. 8-11.

The amount of precession of the antenna axis of the satellite 26 isdetermined from the earth by means of the orientation signal developedby the second orientation oscillator 96 associated with the slit 72 andorientation sensing solar cell 73. This determination can be made onlyduring the time of day that the satellite 26 is in sunlight. However, inthe equatorial orbit of the present example, only rarely does the earth100 come between the sun and the satellite 26, and then only forintervals of short duration. As the satellite 26 spins about its spinaxis, the slit 72 periodically passes through sunlight. Thus a periodicorientation signal is developed and transmitted to the earth 100 and theorientation of the antenna axis with respect to the sun may bedetermined. This information tells when the necessary precession hasbeen completed.

ln as much as the satellite 26 may have velocity and altitude errors, ittraverses only an approximate 24- hour stationary or synchronous orbit112, and corrections are made by radio control of the velocity controlvalve 541. The satellite 26 is tracked from the earth 100 by means ofradio signals transmitted to the satellite 26 and relayed back to theearth 100 to determine the drift of the satellite 26 relative to theearth 100. The velocity of the satellite 26 is increased or decreased byopening the velocity control valve 54 for controlled time intervalsduring the proper portion of the spin cycle by means of radio controlsignals. Opening of the velocity control valve 54 results in jets ofcompressed nitrogen issuing from the associated nozzle 56 to provide areactive force that changes the velocity of the satellite 26. This willcorrect the orbital period and will also reduce or eliminate theeccentricity of the orbit. Inclination of the orbit may be corrected bytiming the attitude control valve 55 and nozzle 57 so that on periods ofthe jet produce no net precession. For example, the valve 55 may beopened for one complete revolution of the satellite 26 about the spinaxis, which produces no net torque about a single axis and hence no netprecession.

As stated hereinbefore, the antenna axis of the satellite 26 is the axisof the maximum moment of inertia. This provides stability against theeffects of vibration and associated energy loss that would otherwisetend to orient the spin around the axis of the largest moment ofinertia, if it were other than the antenna axis. In this way, theeffects of such vibration are to cause the spin to stabilize about thedesired axis, that is, to damp the nutation. In addition to the naturaltendency of the satellite 26 to damp nutation, the nutation damper 58further reduces any nutational motion by absorbing the nutation energy.The nutation damper 58 is resonant at the correct frequency regardlessof spin if the radius of curvature of the tube 60 is chosen as indicatedhereinbefore. When precession is accomplished in a constant direction bya series of pulses, the nutation resulting from a single pulse does notbuild up so that only a small damper S8 is required.

Thus, the satellite 26 is made to orbit around the earth 100 with thesame angular velocity as the earth 100 and in the same direction ofrotation. The antenna axis of the satellite 26 is parallel to the axisof rotation 1102 of the earth 100 so that the antenna 50 radiatessignals to the earth 100 in a narrow beam and the solar cells 52 areoriented to receive optimum light from the sun.

As to the disclosed means for sensing the orientation of the satellite26, namely, the slits 70 and 72, orientation sensing solar cells 71 and73, and the orientation oscillators 95 and 96, it will be understoodthat other means may be provided. For example, asymmetry may bedeliberately introduced into the antenna radiation pattern ofthesatellite 26.

It will be obvious that more than one jet nozzle may be provided forvelocity control or for attitude control. Several jets of gas may beproduced at different locations on a spinning vehicle, each jet beingpulsed and timed or synchronized as described hereinbefore withreference to a single jet. Further, a jet or jets may be produced on aspinning body to achieve simultaneous change in velocity and attitude soas to realize, for example, a proportional navigation course. By thismeans a target seeking vehicle may be constructed in accordance with theinvention.

It will also be apparent that valves may be pulsed or modulated inaccordance with desired functions of time, such as portions of sinewaves, for example, to provide smoother control. In addition, jetcontrol systems in accordance with the invention may be used for othertypes of satellites than communication satellites. For example,meterological, astronomical or navigational satellites may also bespin-stabilized and controlled by pulsed jets of gas to change velocityor attitude of the spin axis in space, or both. Furthermore, theprinciples embodied in the present invention may also be applied tovehicles for probing into space and for intercepting other spacevehicles.

Thus, there has been described a method and apparatus for launching asatellite into a particular orbit, and with a predetermined orientationwith respect to the earth, to provide optimum antenna gain and optimumsolar cell illumination. By using simple spin-stabilization to orientthe satellite, the weight and complexity of the satellite have beenminimized. A method and apparatus has been described for precessing thespin axis of a spin-stabilized body from a first position to a secondposition perpendicular to the first position, and for damping ofnutation. Further, a method and apparatus has been described forcorrecting the orbit of a spin-stabilized satellite.

What is claimed is:

l. A nutation damper for a spin-stabilized body comprising: a bodyhaving a spin axis passing through the center of gravity thereof;arcuate containing means disposed within said body at a distance fromsaid axis and arranged substantially parallel thereto; the curvature ofsaid containing means being concave toward said axis; the center of saidcontaining means being on a line normal to said axis passing through thecenter of gravity of said body; means providing viscosity disposedwithin said containing means; and a movable weight disposed within saidcontaining means.

2. A nutation damper for a spin-stabilized body comprising: a bodyhaving a spin axis passing through the center of gravity thereof; anarcuate tube disposed within said body at a distance from said spin axisand arranged substantially parallel to said spin axis; the curvature ofsaid tube being concave toward said spin axis; the center of said tubebeing on a line normal to said spin axis passing through the center ofgravity of said body, the radius of curvature of said tube being equalto the distance of said tube from said spin axis divided by 1.4 andmultiplied by (I /I, l where I, is equal to the moment of inertia ofsaid body around any axis normal to said spin axis passing through thecenter of gravity of said body and I, is the moment of inertia of saidbody around said spin axis; a viscous fluid sealed within said tube; anda ball disposed within said tube.

3. A nutation damper for use in a spin-stabilized body having a spinaxis passing through the center of gravity thereof comprising: elongatedcontaining means disposed within said body at a distance from said axisand arranged substantially parallel thereto; the center of saidcontaining means being on a line normal to said axis and passing throughthe center of gravity of said body; a mass disposed within saidcontaining means and adapted to move along said containing means inresponse to nutation of said body whereby to damp said nutation.

4. A nutation damper for use with a spin-stabilized body having a spinaxis passing through the center of gravity thereof comprising: arcuatecontaining means for attachment to said body at a distance from saidaxis and substantially parallel thereto; said containing means to bedisposed with the curvature thereof being concave toward said axis; thecenter of aid containing means to be disposed on a line normal to saidaxis and passing through the center of gravity of said body; meansproviding viscosity disposed within said containing means; and a massdisposed within said containing means and adapted to move along saidcontaining means in response to nutation of said body whereby to dampsaid nutation.

5. A nutation damper for use with a spin-stabilized body having a spinaxis passing through the center of gravity thereof comprising: arcuatecontaining means for attachment to said body at a distance from saidaxis and substantially parallel thereto; said containing means to bedisposed with the curvature thereof being concave toward said axis; thecenter of said containing means to be disposed on a line normal to saidaxis and passing through the center of gravity of said body; the radiusof curvature of said containing means to be substantially equal to thedistance of said containing means from said axis divided by 1.4 andmultiplied by (I /I I where I, is equal to the moment of inertia of saidbody around any axis normal to said spin axis and I is the moment ofinertia of said body around said spin axis; means providing viscositydisposed within said containing means; and a mass disposed within saidcontaining means and adapted to move along said containing means inresponse to nutation of said body whereby to damp said nutation.

6. A nutation damper for use with a spin-stabilized body having a spinaxis passing through the center of gravity thereof comprising: elongatedcontaining means for attachment to said body at a distance from saidaxis and substantially parallel thereto; the center of I said containingmeans to be disposed on a line normal to said axis and passing throughthe center of gravity of said body; and a mass disposed within saidcontaining means and adapted to move along said containing means inresponse to nutation of said body whereby to damp said nutation.

7. In spin-stabilized apparatus of the class wherein a spin-stabilizedbody having a spin axis passing through the center of gravity thereofhas nutation induced in the motion of the body through the action ofpulsed jet thrust means, the combination with said body ofa nutationdamper having elongated containing means for attachment to said body ata distance from said axis and substantially parallel thereto; the centerof said containing means to be disposed on a line normal to said axisand passing through the center of gravity of said body; and a massdisposed within said containing means and adapted to move along saidcontaining means in response to nutation of said body whereby to dampsaid nutation.

1. A nutation damper for a spin-stabilized body comprising: a bodyhaving a spin axis passing through the center of gravity thereof;arcuate containing means disposed within said body at a distance fromsaid axis and arranged substantially parallel thereto; the curvature ofsaid containing means being concave toward said axis; the center of saidcontaining means being on a line normal to said axis passing through thecenter of gravity of said body; means providing viscosity disposedwithin said containing means; and a movable weight disposed within saidcontaining means.
 2. A nutation damper for a spin-stabilized bodycomprising: a body having a spin axis passing through the center ofgravity thereof; an arcuate tube disposed within said body at a distancefrom said spin axis and arranged substantially parallel to said spinaxis; the curvature of said tube being concave toward said spin axis;the center of said tube being on a line normal to said spin axis passingthrough the center of gravity of said body, the radius of curvature ofsaid tube being equal to the distance of said tube from said spin axisdivided by 1.4 and multiplied by (Ix/Iz - Ix)2; where Ix is equal to themoment of inertia of said body around any axis normal to said spin axispassing through the center of gravity of said body and Iz is the momentof inertia of said body around said spin axis; a viscous fluid sealedwithin said tube; and a ball disposed within said tube.
 3. A nutationdamper for use in a spin-stabilized body having a spin axis passingthrough the center of gravity thereof comprising: elongated containingmeans disposed within said body at a distance from said axis andarranged substantially parallel thereto; the center of said containingmeans being on a line normal to said axis and passing through the centerof gravity of said body; a mass disposed within said containing meansand adapted to move along said containing means in response to nutationof said body whereby to damp said nutation.
 4. A nutation damper for usewith a spin-stabilized body having a spin axis passing through thecenter of gravity thereof comprising: arcuate containing means forattachment to said body at a distance from said axis and substantiallyparallel thereto; said containing means to be disposed with thecurvature thereof being concave toward said axis; the center of aidcontaining means to be disposed on a line normal To said axis andpassing through the center of gravity of said body; means providingviscosity disposed within said containing means; and a mass disposedwithin said containing means and adapted to move along said containingmeans in response to nutation of said body whereby to damp saidnutation.
 5. A nutation damper for use with a spin-stabilized bodyhaving a spin axis passing through the center of gravity thereofcomprising: arcuate containing means for attachment to said body at adistance from said axis and substantially parallel thereto; saidcontaining means to be disposed with the curvature thereof being concavetoward said axis; the center of said containing means to be disposed ona line normal to said axis and passing through the center of gravity ofsaid body; the radius of curvature of said containing means to besubstantially equal to the distance of said containing means from saidaxis divided by 1.4 and multiplied by (Ix/Iz - Ix)2; where Ix is equalto the moment of inertia of said body around any axis normal to saidspin axis and Iz is the moment of inertia of said body around said spinaxis; means providing viscosity disposed within said containing means;and a mass disposed within said containing means and adapted to movealong said containing means in response to nutation of said body wherebyto damp said nutation.
 6. A nutation damper for use with aspin-stabilized body having a spin axis passing through the center ofgravity thereof comprising: elongated containing means for attachment tosaid body at a distance from said axis and substantially parallelthereto; the center of said containing means to be disposed on a linenormal to said axis and passing through the center of gravity of saidbody; and a mass disposed within said containing means and adapted tomove along said containing means in response to nutation of said bodywhereby to damp said nutation.
 7. In spin-stabilized apparatus of theclass wherein a spin-stabilized body having a spin axis passing throughthe center of gravity thereof has nutation induced in the motion of thebody through the action of pulsed jet thrust means, the combination withsaid body of a nutation damper having elongated containing means forattachment to said body at a distance from said axis and substantiallyparallel thereto; the center of said containing means to be disposed ona line normal to said axis and passing through the center of gravity ofsaid body; and a mass disposed within said containing means and adaptedto move along said containing means in response to nutation of said bodywhereby to damp said nutation.